Gas turbine engine component with converging/diverging cooling passage

ABSTRACT

A component for a gas turbine engine includes a gas path wall having a first surface and a second surface and a cooling hole extending through the gas path wall from the first surface to the second surface. The cooling hole includes an inlet portion having an inlet at the first surface, an outlet portion having an outlet at the second surface, and a transition defined between the inlet and the outlet. The inlet portion converges in a first direction from the inlet to the transition and diverges in a second direction from the inlet to the transition. The outlet portion diverges at least in one of the first and second directions from the transition to the outlet.

CROSS REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Application No.61/599,359, filed on Feb. 15, 2012, and entitled “Gas Turbine EngineComponent with Converging/Diverging Cooling Passage,” the disclosure ofwhich is incorporated by reference in its entirety.

BACKGROUND

This invention relates generally to turbomachinery, and specifically toturbine flow path components for gas turbine engines. In particular, theinvention relates to cooling techniques for airfoils and other gasturbine engine components exposed to hot working fluid flow, including,but not limited to, rotor blades and stator vane airfoils, endwallsurfaces including platforms, shrouds and compressor and turbinecasings, combustor liners, turbine exhaust assemblies, thrust augmentorsand exhaust nozzles.

Gas turbine engines are rotary-type combustion turbine engines builtaround a power core made up of a compressor, combustor and turbine,arranged in flow series with an upstream inlet and downstream exhaust.The compressor section compresses air from the inlet, which is mixedwith fuel in the combustor and ignited to generate hot combustion gas.The turbine section extracts energy from the expanding combustion gas,and drives the compressor section via a common shaft. Expandedcombustion products are exhausted downstream, and energy is delivered inthe form of rotational energy in the shaft, reactive thrust from theexhaust, or both.

Gas turbine engines provide efficient, reliable power for a wide rangeof applications in aviation, transportation and industrial powergeneration. Small-scale gas turbine engines typically utilize aone-spool design, with co-rotating compressor and turbine sections.Larger-scale combustion turbines including jet engines and industrialgas turbines (IGTs) are generally arranged into a number of coaxiallynested spools. The spools operate at different pressures, temperaturesand spool speeds, and may rotate in different directions.

Individual compressor and turbine sections in each spool may also besubdivided into a number of stages, formed of alternating rows of rotorblade and stator vane airfoils. The airfoils are shaped to turn,accelerate and compress the working fluid flow, or to generate lift forconversion to rotational energy in the turbine.

Industrial gas turbines often utilize complex nested spoolconfigurations, and deliver power via an output shaft coupled to anelectrical generator or other load, typically using an external gearbox.In combined cycle gas turbines (CCGTs), a steam turbine or othersecondary system is used to extract additional energy from the exhaust,improving thermodynamic efficiency. Gas turbine engines are also used inmarine and land-based applications, including naval vessels, trains andarmored vehicles, and in smaller-scale applications such as auxiliarypower units.

Aviation applications include turbojet, turbofan, turboprop andturboshaft engine designs. In turbojet engines, thrust is generatedprimarily from the exhaust. Modern fixed-wing aircraft generally employturbofan and turboprop configurations, in which the low pressure spoolis coupled to a propulsion fan or propeller. Turboshaft engines areemployed on rotary-wing aircraft, including helicopters, typically usinga reduction gearbox to control blade speed. Unducted (open rotor)turbofans and ducted propeller engines also known, in a variety ofsingle-rotor and contra-rotating designs with both forward and aftmounting configurations.

Aviation turbines generally utilize two and three-spool configurations,with a corresponding number of coaxially rotating turbine and compressorsections. In two-spool designs, the high pressure turbine drives a highpressure compressor, forming the high pressure spool or high spool. Thelow-pressure turbine drives the low spool and fan section, or a shaftfor a rotor or propeller. In three-spool engines, there is also anintermediate pressure spool. Aviation turbines are also used to powerauxiliary devices including electrical generators, hydraulic pumps andelements of the environmental control system, for example using bleedair from the compressor or via an accessory gearbox.

Additional turbine engine applications and turbine engine types includeintercooled, regenerated or recuperated and variable cycle gas turbineengines, and combinations thereof. In particular, these applicationsinclude intercooled turbine engines, for example with a relativelyhigher pressure ratio, regenerated or recuperated gas turbine engines,for example with a relatively lower pressure ratio or for smaller-scaleapplications, and variable cycle gas turbine engines, for example foroperation under a range of flight conditions including subsonic,transonic and supersonic speeds. Combined intercooled andregenerated/recuperated engines are also known, in a variety of spoolconfigurations with traditional and variable cycle modes of operation.

Turbofan engines are commonly divided into high and low bypassconfigurations. High bypass turbofans generate thrust primarily from thefan, which accelerates airflow through a bypass duct oriented around theengine core. This design is common on commercial aircraft andtransports, where noise and fuel efficiency are primary concerns. Thefan rotor may also operate as a first stage compressor, or as apre-compressor stage for the low-pressure compressor or booster module.Variable-area nozzle surfaces can also be deployed to regulate thebypass pressure and improve fan performance, for example during takeoffand landing. Advanced turbofan engines may also utilize a geared fandrive mechanism to provide greater speed control, reducing noise andincreasing engine efficiency, or to increase or decrease specificthrust.

Low bypass turbofans produce proportionally more thrust from the exhaustflow, generating greater specific thrust for use in high-performanceapplications including supersonic jet aircraft. Low bypass turbofanengines may also include variable-area exhaust nozzles and afterburneror augmentor assemblies for flow regulation and short-term thrustenhancement. Specialized high-speed applications include continuouslyafterburning engines and hybrid turbojet/ramjet configurations.

Across these applications, turbine performance depends on the balancebetween higher pressure ratios and core gas path temperatures, whichtend to increase efficiency, and the related effects on service life andreliability due to increased stress and wear. This balance isparticularly relevant to gas turbine engine components in the hotsections of the compressor, combustor, turbine and exhaust sections,where active cooling is required to prevent damage due to high gas pathtemperatures and pressures.

SUMMARY

This invention concerns a component for a gas turbine engine thatincludes a gas path wall having a first surface and a second surface anda cooling hole extending through the gas path wall from the firstsurface to the second surface. The cooling hole includes an inletportion having an inlet at the first surface, an outlet portion havingan outlet at the second surface, and a transition defined between theinlet and the outlet. The inlet portion converges in a first directionfrom the inlet to the transition and diverges in a second direction fromthe inlet to the transition. The outlet portion diverges at least in oneof the first and second directions from the transition to the outlet.

Another embodiment of the present invention is an airfoil that includesa wall having a first surface and a second surface that is exposed tohot working fluid flow. A cooling hole includes a metering sectionhaving an inlet at the first surface, a diffusing section having anoutlet at the second surface, and a transition defined between the inletand the outlet. The metering section converges in a first direction fromthe inlet to the transition, and diverges in a second direction from theinlet to the transition. The diffusing section diverges at least in oneof the first and second directions from the transition to the outlet.

Another embodiment of the present invention is a gas turbine enginecomponent that includes a gas path wall having a first surface and asecond surface and a cooling hole extending through the gas path wall.The cooling hole has an inlet portion with an inlet in the firstsurface, an outlet portion with an outlet in the second surface, and atransition defined between the inlet portion and the outlet portion. Afirst cooling hole surface extends along the cooling hole. The firstcooling hole surface is substantially straight from the inlet throughthe transition to the outlet. A second cooling hole surface extendingalong the cooling hole opposite the first cooling hole surface. Thesecond cooling hole surface converges toward the first cooling holesurface from the inlet to the transition and diverges away from thefirst cooling hole surface from the transition to the outlet.

Another embodiment of the present invention is a component for a gasturbine engine. The component includes a flow path wall having a firstsurface and a second surface. The first surface is exposed to coolingfluid and the second surface is exposed to hot working fluid. A coolinghole includes a metering section having an inlet at the first surface, adiffusing section having an outlet at the second surface, and atransition defined between the inlet and the outlet. An upstream surfaceof the cooling hole is substantially straight from the inlet to theoutlet. A downstream surface of the cooling hole converges toward theupstream surface from the inlet to the transition and diverges away fromthe upstream surface from the transition to the outlet.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional view of a gas turbine engine.

FIG. 2A is a perspective view of an airfoil for the gas turbine engine,in a rotor blade configuration.

FIG. 2B is a perspective view of an airfoil for the gas turbine engine,in a stator vane configuration.

FIG. 3A is a cross-sectional view of the gas path wall for a cooled gasturbine engine component, taken in a longitudinal direction.

FIG. 3B is an alternate cross-sectional view of the gas path wall,showing the cooling hole geometry.

FIG. 3C is a transverse cross-sectional view of the gas path wall.

FIG. 4A is a schematic view of the gas path wall, illustrating thecooling hole geometry.

FIG. 4B is a schematic view of the gas path wall, illustrating analternate cooling hole geometry in the inlet region.

FIG. 5A is a schematic view of the gas path wall, illustrating thecooling hole geometry in the outlet region.

FIG. 5B is a schematic view of the gas path wall, illustrating analternate cooling hole geometry in the outlet region.

FIG. 6 is a block diagram of a method for forming a cooling hole in agas turbine engine component.

DETAILED DESCRIPTION

FIG. 1 is a cross-sectional view of gas turbine engine 10. Gas turbineengine (or turbine engine) 10 includes a power core with compressorsection 12, combustor 14 and turbine section 16 arranged in flow seriesbetween upstream inlet 18 and downstream exhaust 20. Compressor section12 and turbine section 16 are arranged into a number of alternatingstages of rotor airfoils (or blades) 22 and stator airfoils (or vanes)24.

In the turbofan configuration of FIG. 1, propulsion fan 26 is positionedin bypass duct 28, which is coaxially oriented about the engine corealong centerline (or turbine axis) C_(L). An open-rotor propulsion stage26 may also provided, with turbine engine 10 operating as a turboprop orunducted turbofan engine. Alternatively, fan rotor 26 and bypass duct 28may be absent, with turbine engine 10 configured as a turbojet orturboshaft engine, or an industrial gas turbine.

For improved service life and reliability, components of gas turbineengine 10 are provided with an improved cooling configuration, asdescribed below. Suitable components for the cooling configurationinclude rotor airfoils 22, stator airfoils 24 and other gas turbineengine components exposed to hot gas flow, including, but not limitedto, platforms, shrouds, casings and other endwall surfaces in hotsections of compressor 12 and turbine 16, and liners, nozzles,afterburners, augmentors and other gas wall components in combustor 14and exhaust section 20.

In the two-spool, high bypass configuration of FIG. 1, compressorsection 12 includes low pressure compressor (LPC) 30 and high pressurecompressor (HPC) 32, and turbine section 16 includes high pressureturbine (HPT) 34 and low pressure turbine (LPT) 36. Low pressurecompressor 30 is rotationally coupled to low pressure turbine 36 via lowpressure (LP) shaft 38, forming the LP spool or low spool. High pressurecompressor 32 is rotationally coupled to high pressure turbine 34 viahigh pressure (HP) shaft 40, forming the HP spool or high spool.

Flow F at inlet 18 divides into primary (core) flow F_(P) and secondary(bypass) flow F_(S) downstream of fan rotor 26. Fan rotor 26 acceleratessecondary flow F_(S) through bypass duct 28, with fan exit guide vanes(FEGVs) 42 to reduce swirl and improve thrust performance. In somedesigns, structural guide vanes (SGVs) 42 are used, providing combinedflow turning and load bearing capabilities.

Primary flow F_(P) is compressed in low pressure compressor 30 and highpressure compressor 32, then mixed with fuel in combustor 14 and ignitedto generate hot combustion gas. The combustion gas expands to providerotational energy in high pressure turbine 34 and low pressure turbine36, driving high pressure compressor 32 and low pressure compressor 30,respectively. Expanded combustion gases exit through exhaust section (orexhaust nozzle) 20, which can be shaped or actuated to regulate theexhaust flow and improve thrust performance.

Low pressure shaft 38 and high pressure shaft 40 are mounted coaxiallyabout centerline C_(L), and rotate at different speeds. Fan rotor (orother propulsion stage) 26 is rotationally coupled to low pressure shaft38. In advanced designs, fan drive gear system 44 is provided foradditional fan speed control, improving thrust performance andefficiency with reduced noise output.

Fan rotor 26 may also function as a first-stage compressor for gasturbine engine 10, and LPC 30 may be configured as an intermediatecompressor or booster. Alternatively, propulsion stage 26 has an openrotor design, or is absent, as described above. Gas turbine engine 10thus encompasses a wide range of different shaft, spool and turbineengine configurations, including one, two and three-spool turboprop and(high or low bypass) turbofan engines, turboshaft engines, turbojetengines, and multi-spool industrial gas turbines.

In each of these applications, turbine efficiency and performance dependon the overall pressure ratio, defined by the total pressure at inlet 18as compared to the exit pressure of compressor section 12, for exampleat the outlet of high pressure compressor 32, entering combustor 14.Higher pressure ratios, however, also result in greater gas pathtemperatures, increasing the cooling loads on rotor airfoils 22, statorairfoils 24 and other components of gas turbine engine 10. To reduceoperating temperatures, increase service life and maintain engineefficiency, these components are provided with improved coolingconfigurations, as described below. Suitable components include, but arenot limited to, cooled gas turbine engine components in compressorsections 30 and 32, combustor 14, turbine sections 34 and 36, andexhaust section 20 of gas turbine engine 10.

FIG. 2A is a perspective view of rotor airfoil (or blade) 22 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Rotor airfoil 22 extends axially from leading edge 51 to trailing edge52, defining pressure surface 53 (front) and suction surface 54 (back)therebetween.

Pressure and suction surfaces 53 and 54 form the major opposing surfacesor walls of airfoil 22, extending axially between leading edge 51 andtrailing edge 52, and radially from root section 55, adjacent innerdiameter (ID) platform 56, to tip section 57, opposite ID platform 56.In some designs, tip section 57 is shrouded.

Cooling holes or outlets 60 are provided on one or more surfaces ofairfoil 22, for example along leading edge 51, trailing edge 52,pressure (or concave) surface 53, or suction (or convex) surface 54, ora combination thereof. Cooling holes or passages 60 may also be providedon the endwall surfaces of airfoil 22, for example along ID platform 56,or on a shroud or engine casing adjacent tip section 57.

FIG. 2B is a perspective view of stator airfoil (or vane) 24 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Stator airfoil 24 extends axially from leading edge 61 to trailing edge62, defining pressure surface 63 (front) and suction surface 64 (back)therebetween. Pressure and suction surfaces 63 and 64 extend from inner(or root) section 65, adjacent ID platform 66, to outer (or tip) section67, adjacent outer diameter (OD) platform 68.

Cooling holes or outlets 60 are provided along one or more surfaces ofairfoil 24, for example leading or trailing edge 61 or 62, pressure(concave) or suction (convex) surface 63 or 64, or a combinationthereof. Cooling holes or passages 60 may also be provided on theendwall surfaces of airfoil 24, for example along ID platform 66 and ODplatform 68.

Rotor airfoils 22 (FIG. 2A) and stator airfoils 24 (FIG. 2B) are formedof high strength, heat resistant materials such as high temperaturealloys and superalloys, and are provided with thermal anderosion-resistant coatings. Airfoils 22 and 24 are also provided withinternal cooling passages and cooling holes 60 to reduce thermal fatigueand wear, and to prevent melting when exposed to hot gas flow in thehigher temperature regions of a gas turbine engine or otherturbomachine. Cooling holes 60 deliver cooling fluid (e.g., steam or airfrom a compressor) through the outer walls and platform structures ofairfoils 22 and 24, creating a thin layer (or film) of cooling fluid toprotect the outer (gas path) surfaces from high temperature flow.

While surface cooling extends service life and increases reliability,injecting cooling fluid into the gas path also reduces engineefficiency, and the cost in efficiency increases with the requiredcooling flow. Cooling holes 60 are thus provided with improved meteringand inlet geometry to reduce jets and blow off, and improved diffusionand exit geometry to reduce flow separation and corner effects. Coolingholes 60 reduce flow requirements and improve the spread of coolingfluid across the hot surfaces of airfoils 22 and 24, and other gasturbine engine components, so that less flow is needed for cooling andefficiency is maintained or increased.

FIG. 3A is a cross-sectional view of gas turbine engine component(turbine or turbomachinery component) 100 with gas path wall 102, takenin a longitudinal direction and that carries a cool first surface 106and an opposite, hot, second surface 108. Cooling hole 104 extendsthrough gas path wall 102 from first surface 106 to second surface 108to form cooling hole 60 in, for example the outer wall of an airfoil,casing, combustor liner, exhaust nozzle or other gas turbine enginecomponent, as described above.

Gas path wall 102 of component 100 is exposed to cooling fluid on firstsurface 106 with longitudinal hot gas or working fluid flow H alongsecond surface 108. In some components, for example airfoils, firstsurface 106 is an inner surface (or inner wall) and second surface 108is an outer surface (or outer wall). In other components, for examplecombustor liners and exhaust nozzles, first surface 106 is an outersurface (or outer wall), and second surface 108 is an inner surface (orinner wall). More generally, the terms inner and outer are merelyrepresentative, and may be interchanged.

Cooling hole 104 delivers cooling fluid C from first surface 106 of wall102 to second surface 108, for example to provide diffusive flow andfilm cooling. Cooling hole 104 is also inclined along axis A in adownstream direction, in order to improve cooling fluid coverage oversecond surface 108, with less separation and reduced flow mixing.

Axis A is an approximate longitudinal axis of flow of metering section110. Cooling hole 104 includes metering section 110 and diffusingsection 112, and extends along axis A from metering section 110 todiffusing section 112. Metering section 110 has inlet 114 at firstsurface 106 of gas path wall 102, and diffusing section 112 has outlet116 at second surface 108 of gas path wall 102. Outlet 116 defines aperimeter of diffusing section 112 at an intersection of diffusingsection 112 and second surface 108. Surfaces 120, 122, 130, and 132 ofcooling hole 104 define cooling hole 104 between inlet 114 and outlet116.

Transition 118 is defined in the region between metering section 110 anddiffusing section 112, where cooling hole 104 becomes divergent(increasing flow area) and where the cooling fluid flow becomesdiffusive. Transition 118 may be relatively abrupt or may encompass anextended portion of cooling hole 104, for example in a flow transitionregion between metering section 110 and diffusing section 112, or over aregion of overlap between metering section 110 and diffusing section112.

As shown in FIG. 3A, metering section 110 of cooling hole 104 convergesalong axis A between inlet 114 and transition 118, as defined betweenopposite upstream and downstream boundaries or surfaces 120 and 122. Inparticular, upstream surface 120 and downstream surface 122 convergetoward one another in the longitudinal direction, in the region frominlet 114 through metering section 110 to transition 118. This decreasesthe cross sectional area (or flow area) of metering section 110, inorder to regulate the cooling fluid flow between inlet 114 and outlet116. Though surfaces 120 and 122 are represented in the cross-sectionalview of FIG. 3A with a line, they can be curved as described furtherbelow. In the illustrated embodiment, surfaces 120 and 122 are angledwith respect to both first surface 106 and second surface 108.

Diffusing section 112 of cooling hole 104 diverges between transition118 and outlet 116. That is, upstream and downstream surfaces 120 and122 diverge from one another in the longitudinal direction, in theregion from transition 118 through diffusing section 112 to outlet 116.This increases the lateral width of transition 118, and thus the crosssectional area (or flow area) of diffusing section 112, in order toprovide diffusive flow between transition 118 and outlet 116.

FIG. 3B is an alternate cross-sectional view of gas path wall 102,showing a different geometry for cooling hole 104. In this design,upstream surface 120 is substantially linear from inlet 114 throughtransition 118 to outlet 116, in order to decrease flow separation intransition region 118. As shown in the FIG. 3B, axis A of passage 104may be taken substantially parallel to upstream surface 120.

Regardless of the upstream wall configuration, one or more longitudinalridges 124 may be formed on downstream surface 122 of diffusing section112. Longitudinal ridges 124 project out from downstream surface 122 toseparate diffusing section 112 of cooling hole 104 into different lobes,discouraging swirl and reducing flow separation as described below withrespect to FIGS. 5A and 5B.

Longitudinal ridges 124 form ribs or ridge structures extending alongdownstream surface 122, for example from transition 118 toward trailingedge (downstream end) 126 of outlet 116, as shown in FIG. 3B. Transitionregion 128 can extend from longitudinal ridge 124 to trailing edge 126of outlet 116, in order to reduce separation and improve downstreamcooling performance. Transition region 128 may be formed along thesurface of downstream surface 122, as described below. Transition region128 can be flat or planar. Alternatively, transition region 128 can benon-flat and non-planar, such as curved (e.g. convex) longitudinallyand/or laterally.

FIG. 3C is a transverse cross sectional view of gas path wall 102, takenalong axis A and looking in a downstream direction, in a planeperpendicular or transverse to the longitudinal cross sections of FIGS.3A and 3B. In this downstream view, hot gas flow H is directed into thepage, and lateral side surfaces 130 and 132 are separated in thetransverse direction across axis A, perpendicular to hot gas flow H.

As shown in the FIG. 3C, metering section 110 of cooling hole 104 isdivergent in the lateral or transverse direction, and first and secondside walls (or lateral boundaries) 130 and 132 diverge from one another(and from axis A) between inlet 114 and transition 118. While meteringsection 110 is convergent in the longitudinal direction, therefore, asshown in FIG. 3A, metering section 110 is divergent in the transversedirection, as shown in FIG. 3C. Alternatively, metering section 110 canbe convergent in the transverse direction and/or divergent in thelongitudinal direction, as shown in FIG. 4B.

In metering section 110, first side surface 130 diverges from secondside surface 132 at a rate that complements the rate of convergence ofupstream and downstream surfaces 120 and 122 (shown in FIGS. 3A and 3B)so that the flow area of metering section 110 either remains constant ordecreases from inlet 114 to transition 118. Thus, metering section 110restricts flow from inlet 114 through transition 118, regulating theflow rate and improving efficiency by providing only the desired levelof cooling fluid flow to diffusing section 112, for more efficientcooling and greater coverage along second surface 108 of gas path wall102, downstream of outlet 116.

In diffusing section 112, side surfaces 130 and 132 diverge laterallyfrom one another (and from axis A) between transition 118 and outlet116. Thus, diffusing section 112 is divergent in both the longitudinaldirection of FIG. 3A, and in the transverse direction of FIG. 3C. Thisconfiguration improves diffusive flow between transition 118 and outlet116, decreasing flow separation at trailing edge 126 and improvingcooling performance along second surface 108 of gas path wall 102. Inthe illustrated embodiment, side surfaces 130 and 132 are angled withrespect to both first surface 106 and second surface 108.

FIG. 4A is a schematic view of gas path wall 102, illustrating thegeometry of cooling hole 104 in inlet (metering) portion 110. This is adownward or inward view, looking down on gas path wall 102 and coolinghole 104.

Second surface 108 of gas path wall 102 is exposed to hot gas flow H ina longitudinal and downstream direction, from left to right in FIG. 4A.Cooling hole 104 extends down through gas path wall 102, from outlet 116at second surface 108 (solid lines) to inlet 114 at first surface 106(dashed lines). Transition 118 is between inlet 114 and outlet 116.Inlet or metering section 110 of cooling hole 104 extends from inlet 114to transition 118. Diffusing section 112 of cooling hole 104 extendsfrom transition 118 to outlet 116.

As shown in FIG. 4A, diffusing section 112 of cooling hole 104 diverges(widens) in both the longitudinal direction, along the direction of hotgas flow H, and in the lateral direction, transverse or perpendicular tothe direction of hot gas flow H. That is, the longitudinal and lateraldimensions of diffusing section 112 both increase between transition 118and outlet 116. As a result, the cross sectional (flow) area of coolinghole 104 increases from transition 118 through diffusing section 112 tooutlet 116. This configuration promotes diffusive flow in outlet region112 of cooling hole 104, improving coverage and performance along secondsurface 108 of gas path wall 102, downstream of outlet 116 at trailingedge 126.

Metering section 110, in contrast, converges (narrows) in thelongitudinal direction and diverges (widens) in the lateral direction.That is, the longitudinal dimension of passage 104 decreases betweeninlet 114 and transition 118, while the transverse dimension increases.The relative convergence and divergence are selected so that the crosssectional (flow) area of metering section 110 decreases, or remainssubstantially constant, from inlet 114 through metering section 110 totransition 118.

This configuration promotes regulated flow in inlet region 110 ofcooling hole 104, in order to expand cooling flow area coverage andimprove cooling efficiency. The cross-section of metering section 110also varies and can be selected to improve cooling efficiency, forexample using a circular, elliptical, oblong, or crescent shape, atinlet 114 or transition 118, or a circular, elliptical or oblongcross-sectional geometry between inlet 114 and transition 118.

The configuration of outlet 116 is also selected to improve coolingperformance. In the particular configuration of FIG. 4A, for example,outlet 116 is formed as a delta, with arcuate upstream surface 120extending toward and substantially linear trailing edge 126.Alternatively, trailing edge 126 may be convex.

FIG. 4B is a schematic view of gas path wall 102, illustrating analternate geometry for metering section 110 of cooling hole 104. Asshown in FIG. 3A, the transverse dimension of metering section 110decreases between inlet 114 and transition 118, while the lateraldimension can increase. Thus, metering section 110 of cooling hole 104diverges in the longitudinal direction, and converges in the transversedirection. Again, the relative convergence and divergence can beselected so that the cross sectional (flow) area of cooling hole 104decreases, or remains substantially constant, from inlet 114 throughmetering section 110 to transition 118.

FIG. 5A is a schematic view of gas path wall 102, illustrating thegeometry of cooling hole 104 in outlet (diffusion) portion 112. Thegeometry of outlet 116 is selected to reduce flow separation,particularly along trailing edge 126. In particular, trailing edge 126is substantially straight or convex in the downstream sense, in orderincrease laminar flow and decrease flow separation along outer gas pathwall surface 108.

FIG. 5A also shows longitudinal ridges 124 in diffusing section 112 ofcooling hole 104, extending from transition 118 to trailing edge 126 ofoutlet 116. Longitudinal ridges 124 project out (upward) from downstreamsurface 122 of cooling hole 104, dividing diffusing section 112 intomultiple lobes 136 and discouraging transverse (swirl or vortex) flowcomponents. This configuration reduces cooling fluid losses andseparation at outlet 116, for improved coverage and cooling efficiencyalong second surface 108 of gas path wall 102, downstream from trailingedge 126 of outlet 116.

For example, longitudinal ridges 124 may be formed at the intersectionor interface between adjacent lobes 136, where lobes 136 have arcuate orcurved surfaces along downstream surface 122, meeting at convex orcusp-shaped longitudinal ridges 124. Alternatively, longitudinal ridges124 may be formed at the intersection or interface between adjacentlobes 136 with substantially planar surfaces along downstream surface122, meeting at a triangular ridge or intersection between planes ofdifferent slope in adjacent lobes 136. Lobes 136 are surfaces of wall102 which define distinct channel-like portions of the void of coolinghole 104 at diffusing section 112. In these designs, longitudinal ridges124 diverge in the direction from transition 118 toward trailing edge126 of outlet 116, between diverging (diffusion) segments 134 of lateralside surfaces 130 and 132.

The geometries of longitudinal ridge or divider processes 124 thus vary.Longitudinal ridges 124 may also be formed as long, narrow featuresextending along the wall of cooling hole 104, for example where twosloping sides (e.g., of lobes 136) meet, or as a narrow raised band orrib structure. Longitudinal ridges 124 may also be either substantiallypointed or rounded, for example where two curved lobe or wall surfacesmeet, or where the direction of curvature reverses along the wall ofcooling hole 104. Longitudinal ridges 124 may also be formed as archedor cone-shape features, for example at the boundary of two lobes 136.

FIG. 5B is a schematic view of gas path wall 102, illustrating analternate geometry for diffusing section 112 of cooling hole 104. Inthis configuration, longitudinal ridges 124 extend from transition 118to terminate at planar transition region 128. Planar transition region128 is without longitudinal ridges 124, and coextends with downstreamsurface 122 of cooling hole 104, from the downstream end of longitudinalridges 124 to trailing edge 126 of outlet 116.

Transition region 128 of outlet portion 112 eliminates cusps alongtrailing edge 126, further reducing flow separation for improved coolingperformance along second surface 108 of gas path wall 102. Thedimensions of longitudinal ridges 124 are selected to discourage swirl,as described above, and the dimensions of transition region 128 areselected to improve flow uniformity along trailing edge 126.

The overall geometry of cooling hole 104 thus varies, as describedabove, and as shown in the figures. The design of inlet 114 and outlet116 may also vary, including various circular, oblate, oval, crescent,trapezoidal, triangular, cusped and delta shaped profiles with arcuate,curved, angular or piecewise linear upstream surfaces 120 extending tostraight or convex trailing edges 126. The configuration of cooling hole104 is not limited to these particular examples, moreover, but alsoencompasses different combinations of the various features that areshown, including metering sections 110 and transitions 118 with avariety of different circular, elliptical, oblong and cusped crosssections, and diffusing sections 112 with one, two or three lobes 136,in combination with different transition regions 128.

The cross sectional shape of cooling hole 104 may also vary, for examplealong metering section 110 between inlet 114 and transition 118, withintransition 118, or along diffusing section 112 between transition 118and outlet 116. In particular, different portions of metering section110 may have circular, longitudinal oval, or transverse oval crosssections, as taken across axis A, or an oblong or delta shape, forexamples as shown for outlet 116.

FIG. 6 is a block diagram illustrating method 200 for forming a coolingflow passage through the gas path wall of a gas turbine enginecomponent. For example, method 200 may be used to form cooling hole 60or cooling hole 104 in an airfoil, casing, liner, combustor, augmentoror turbine exhaust component, as described above. Method 200 includesforming a cooling hole in a gas path wall of the component (step 202),forming a metering section in the cooling hole (step 204), and forming adiffusing section adjacent the metering section (step 206).

Forming a cooling hole in the gas path wall (step 202) includes formingan inlet (step 208) in a first surface of the gas path wall, andextending the cooling hole through a transition (step 210) to an outlet(step 212) in the second surface of the gas path wall. These steps maybe performed in any order, depending on component geometry, for exampleby forming the passage starting from the outlet in the outer wall (step212), through the transition (step 210) and to the inlet in the innerwall (step 208). Alternatively, the inlet and outlet may be formed(steps 208 and 212) in a single drilling or manufacturing process, andthe transition may be formed while extending the metering section (step204) from the inlet, or while extending the diffusing section (step 206)from the outlet, as described below.

Forming a metering section (step 204) includes extending the coolinghole from the inlet (at the first surface) to the transition. Themetering section is convergent in a first direction from the inlet tothe transition, and divergent in a second direction from the inlet tothe transition. The convergence and divergence are selected so that theflow area of the passage decreases or remains substantially constantfrom the inlet through the metering section to the outlet, in order toregulate the cooling fluid flow for improved efficiency. Thisconvergence and divergence of the metering section can be formed byfirst laser drilling the metering section and then coated. The coatingcan be deposited in the metering section such that certain portions ofthe metering section receive more coating deposition than other portionsof the metering section. Thus, the metering section can converge in afirst direction and diverge in a second direction.

The first and second directions may be longitudinal or transverse withrespect to hot gas flow along the second surface of the gas path wall.In addition, the upstream wall of the cooling hole can be substantiallystraight from the inlet through the transition to the outlet. Thedownstream wall of the cooling hole may converge toward the upstreamwall, in the region from the inlet through the metering section to thetransition.

Forming a diffusing section (step 206) includes extending the coolinghole from the transition to the outlet (at the second surface). Thediffusing section is convergent in both the first and second directions.Thus, the flow area of the cooling hole increases from the transitionthrough the diffusing section to the outlet, in order to improve coolingfluid distribution and cooling performance. In some methods,longitudinal features are formed (step 214) in the diffusing section, asdescribed above, dividing the outlet portion of the flow passage intolobes in order to discourage swirl and reduce losses.

The gas turbine engine components, gas path walls and cooling holesdescribed herein can thus be manufactured using one or more of a varietyof different processes. These techniques provide each cooling hole withits own particular configuration and features, including, but notlimited to, inlet, metering, transition, diffusion, outlet, upstreamsurface, downstream surface, lateral surface, longitudinal, lobe anddownstream edge features, as described above. In some cases, multipletechniques can be combined to improve overall cooling performance orreproducibility, or to reduce manufacturing costs.

Suitable manufacturing techniques for forming the cooling configurationsdescribed here include, but are not limited to, electrical dischargemachining (EDM), laser drilling, laser machining, electrical chemicalmachining (ECM), water jet machining, casting, conventional machiningand combinations thereof. Electrical discharge machining includes bothmachining using a shaped electrode as well as multiple pass methodsusing a hollow spindle or similar electrode component. Laser machiningmethods include, but are not limited to, material removal by ablation,trepanning and percussion laser machining. Conventional machiningmethods include, but are not limited to, milling, drilling and grinding.

The gas flow path walls and outer surfaces of some gas turbine enginecomponents include one or more coatings, such as bond coats, thermalbarrier coatings, abrasive coatings, abradable coatings and erosion orerosion-resistant coatings. For components having a coating, the inlet,metering section, transition, diffusing section and outlet coolingfeatures may be formed prior to a coating application, after a firstcoating (e.g., a bond coat) is applied, or after a second or third(e.g., interlayer) coating process, or a final coating (e.g.,environmental or thermal barrier) process. Depending on component type,cooling hole or passage location, repair requirements and otherconsiderations, the diffusing section and outlet features may be locatedwithin a wall or substrate, within a thermal barrier coating or othercoating layer applied to a wall or substrate, or combinations thereof.The cooling geometry and other features may remain as described above,regardless of position relative to the wall and coating materials orairfoil materials.

In addition, the order in which cooling features are formed and coatingsare applied may affect selection of manufacturing techniques, includingtechniques used in forming the inlet, metering section, transition,outlet, diffusing section and other cooling features. For example, whena thermal barrier coat or other coating is applied to the second surfaceof a gas path wall before the cooling hole or passage is produced, laserablation or laser drilling may be used. Alternatively, either laserdrilling or water jet machining may be used on a surface without athermal barrier coat. Additionally, different machining methods may bemore or less suitable for forming different features of the coolinghole, for example different laser and other machining techniques may beused for forming the outlet and diffusion features, and for forming thetransition, metering and inlet features.

While the invention is described with reference to exemplaryembodiments, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted withoutdeparting from the spirit and scope of the invention. In addition,different modifications may be made to adapt the teachings of theinvention to particular situations or materials, without departing fromthe essential scope thereof. The invention is thus not limited to theparticular examples disclosed herein, but includes all embodimentsfalling within the scope of the appended claims.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments ofthe present invention.

A component for a gas turbine engine can include a gas path wall havinga first surface and a second surface and a cooling hole extendingthrough the gas path wall from the first surface to the second surface.The cooling hole can include an inlet portion having an inlet at thefirst surface, an outlet portion having an outlet at the second surface,and a transition defined between the inlet and the outlet. The inletportion can converge in a first direction from the inlet to thetransition and diverge in a second direction from the inlet to thetransition. The outlet portion can diverge at least in one of the firstand second directions from the transition to the outlet.

The component of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

the inlet portion of the cooling hole can have an oblong geometry;

cross-sectional area of the cooling hole can decrease or remain constantfrom the inlet to the transition;

cross-sectional area of the cooling hole can increase from thetransition to the outlet;

the first direction can be a longitudinal direction with respect to gasflow along the second surface and the second direction can be atransverse direction with respect to the gas flow along the secondsurface;

the first direction can be a transverse direction with respect to gasflow along the second surface and the second direction can be alongitudinal direction with respect to the gas flow along the secondsurface;

the cooling hole can be inclined in a downstream direction with respectto gas flow along the second surface of the gas path wall;

the cooling hole can have a substantially straight upstream surfaceextending from the inlet through the transition to the outlet;

the cooling hole can have a convergent downstream surface extending fromthe inlet to the transition and a divergent downstream surface extendingfrom the transition to the outlet;

a longitudinal ridge can be formed on a downstream surface of the outletportion between the transition and the outlet, wherein the longitudinalridge divides the outlet portion into lobes;

a transition region can extend between the longitudinal ridge and atrailing edge of the outlet; and/or

the second surface can form a pressure surface, a suction surface or aplatform surface of an airfoil.

An airfoil can include a wall having a first surface and a secondsurface that is exposed to hot working fluid flow. A cooling hole caninclude a metering section having an inlet at the first surface, adiffusing section having an outlet at the second surface, and atransition defined between the inlet and the outlet. The meteringsection can converge in a first direction from the inlet to thetransition, and diverge in a second direction from the inlet to thetransition. The diffusing section can diverge at least in one of thefirst and second directions from the transition to the outlet.

The airfoil of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

the first direction can be a longitudinal direction with respect to thehot working fluid flow and the second direction can be a transversedirection with respect to the hot working fluid flow; and/or

longitudinal ridges can be formed on a downstream wall of the diffusingsection, and the longitudinal ridges can extend between the transitionand the outlet and dividing the diffusing section into lobes.

A gas turbine engine component can include a gas path wall having afirst surface and a second surface and a cooling hole extending throughthe gas path wall. The cooling hole can have an inlet portion with aninlet in the first surface, an outlet portion with an outlet in thesecond surface, and a transition defined between the inlet portion andthe outlet portion. A first cooling hole surface can extend along thecooling hole. The first cooling hole surface can be substantiallystraight from the inlet through the transition to the outlet. A secondcooling hole surface can extend along the cooling hole opposite thefirst cooling hole surface. The second cooling hole surface can convergetoward the first cooling hole surface from the inlet to the transitionand diverge away from the first cooling hole surface from the transitionto the outlet.

The component of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

flow area of the inlet portion can decrease or remain constant betweenthe inlet and the transition, and flow area of the outlet portion canincrease between the transition and the outlet;

the first cooling hole surface can be an upstream surface with respectto hot gas flow along the second surface of the gas path wall and thesecond cooling hole surface can be a downstream surface with respect tothe hot gas flow; and/or

a longitudinal ridge can be formed on the downstream surface between thetransition and the outlet, and the longitudinal ridge can divide theoutlet portion into lobes.

A component for a gas turbine engine can include a flow path wall havinga first surface and a second surface. The first surface can be exposedto cooling fluid and the second surface can be exposed to hot workingfluid. A cooling hole can include a metering section having an inlet atthe first surface, a diffusing section having an outlet at the secondsurface, and a transition defined between the inlet and the outlet. Anupstream surface of the cooling hole can be substantially straight fromthe inlet to the outlet. A downstream surface of the cooling hole canconverge toward the upstream surface from the inlet to the transitionand diverge away from the upstream surface from the transition to theoutlet.

The component of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

first and second lateral side surfaces of the cooling hole can divergefrom the inlet to the transition;

divergence of the first and second lateral side surfaces can be selectedsuch that the cross sectional area of the metering section does notincrease from the inlet to the transition;

divergence of the first and second lateral side surfaces can be selectedsuch that cross sectional area of the metering section decreases fromthe inlet to the transition;

a ridge can extend along the downstream surface of the cooling hole inthe diffusing section, and can divide the diffusing section of thecooling hole into lobes; and/or

a transition region can extend from the ridge to a trailing edge of theoutlet.

The invention claimed is:
 1. A component for a gas turbine engine, thecomponent comprising: a gas path wall having a first surface and asecond surface; and a cooling hole extending through the gas path wallfrom the first surface to the second surface, the cooling holecomprising an inlet portion having an inlet at the first surface, anoutlet portion having an outlet at the second surface, and a transitiondefined between the inlet and the outlet; wherein the inlet portionconverges in a first direction from the inlet to the transition anddiverges in a second direction from the inlet to the transition suchthat a flow area of the inlet portion decreases from the inlet to thetransition or remains substantially constant; and wherein the outletportion diverges at least in one of the first and second directions fromthe transition to the outlet.
 2. The component of claim 1, wherein theinlet portion of the cooling hole has an oblong geometry.
 3. Thecomponent of claim 1, wherein cross-sectional area of the cooling holeincreases from the transition to the outlet.
 4. The component of claim1, wherein the first direction is a longitudinal direction with respectto gas flow along the second surface and the second direction is atransverse direction with respect to the gas flow along the secondsurface.
 5. The component of claim 1, wherein the first direction is atransverse direction with respect to gas flow along the second surfaceand the second direction is a longitudinal direction with respect to thegas flow along the second surface.
 6. The component of claim 1, whereinthe cooling hole is inclined in a downstream direction with respect togas flow along the second surface of the gas path wall.
 7. The componentof claim 6, wherein the cooling hole has a substantially straightupstream surface extending from the inlet through the transition to theoutlet.
 8. The component of claim 7, wherein the cooling hole has aconvergent downstream surface extending from the inlet to the transitionand a divergent downstream surface extending from the transition to theoutlet.
 9. The component of claim 1, further comprising: a longitudinalridge formed on a downstream surface of the outlet portion between thetransition and the outlet, wherein the longitudinal ridge divides theoutlet portion into lobes.
 10. The component of claim 9, furthercomprising: a transition region extending between the longitudinal ridgeand a trailing edge of the outlet.
 11. The component of claim 1, whereinthe second surface forms a pressure surface, a suction surface or aplatform surface of an airfoil.
 12. An airfoil comprising: a wall havinga first surface and a second surface, wherein the second surface isexposed to hot working fluid flow; and a cooling hole comprising ametering section having an inlet at the first surface, a diffusingsection having an outlet at the second surface, and a transition definedbetween the inlet and the outlet; wherein the metering section convergesin a first direction from the inlet to the transition, and diverges in asecond direction from the inlet to the transition such that a flow areaof the metering section decreases from the inlet to the transition orremains substantially constant; and wherein the diffusing sectiondiverges at least in one of the first and second directions from thetransition to the outlet.
 13. The airfoil of claim 12, wherein the firstdirection is a longitudinal direction with respect to the hot workingfluid flow and the second direction is a transverse direction withrespect to the hot working fluid flow.
 14. The airfoil of claim 12,further comprising: longitudinal ridges formed on a downstream wall ofthe diffusing section, the longitudinal ridges extending between thetransition and the outlet and dividing the diffusing section into lobes.15. A gas turbine engine component comprising: a gas path wall having afirst surface and a second surface; a cooling hole extending through thegas path wall, the cooling hole having an inlet portion with an inlet inthe first surface, an outlet portion with an outlet in the secondsurface, and a transition defined between the inlet portion and theoutlet portion; a first cooling hole surface extending along the coolinghole, wherein the first cooling hole surface is substantially straightfrom the inlet through the transition to the outlet; a second coolinghole surface extending along the cooling hole opposite the first coolinghole surface, wherein the second cooling hole surface converges towardthe first cooling hole surface from the inlet to the transition anddiverges away from the first cooling hole surface from the transition tothe outlet; and third and fourth cooling hole surfaces extending alongthe cooling hole, wherein the third and fourth cooling hole surfacesdiverge from one another as the first and second cooling hole surfacesconverge such that a flow area of the inlet portion decreases from theinlet to the transition or remains substantially constant.
 16. The gasturbine engine component of claim 15, wherein the first cooling holesurface is an upstream surface with respect to hot gas flow along thesecond surface of the gas path wall and the second cooling hole surfaceis a downstream surface with respect to the hot gas flow.
 17. The gasturbine engine component of claim 16, further comprising: a longitudinalridge formed on the downstream surface between the transition and theoutlet, wherein the longitudinal ridge divides the outlet portion intolobes.
 18. A component for a gas turbine engine, the componentcomprising: a flow path wall having a first surface and a secondsurface, wherein the first surface is exposed to cooling fluid and thesecond surface is exposed to hot working fluid; a cooling holecomprising a metering section having an inlet at the first surface, adiffusing section having an outlet at the second surface, and atransition defined between the inlet and the outlet; an upstream surfaceof the cooling hole, wherein the upstream surface is substantiallystraight from the inlet to the outlet; and a downstream surface of thecooling hole, wherein the downstream surface converges toward theupstream surface from the inlet to the transition and diverges away fromthe upstream surface from the transition to the outlet; and first andsecond lateral side surfaces of the cooling hole, wherein the first andsecond lateral side surfaces diverge from one another as the upstreamsurface and the downstream surface converge such that a flow area of themetering section decreases from the inlet to the transition or remainssubstantially constant.
 19. The component of claim 18, whereindivergence of the first and second lateral side surfaces is selectedsuch that cross sectional area of the metering section decreases fromthe inlet to the transition.
 20. The component of claim 18, furthercomprising: a ridge extending along the downstream surface of thecooling hole in the diffusing section, wherein the ridge divides thediffusing section of the cooling hole into lobes.
 21. The component ofclaim 20, further comprising: a transition region extending from theridge to a trailing edge of the outlet.